1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade with a serpentine flow cooling circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially in an industrial gas turbine engine, a turbine section includes multiple stages of stator or guide vanes and rotor blades to extract mechanical energy from a hot gas flow passing through the turbine. Increasing the turbine inlet temperature can increase the turbine efficiency, and therefore the engine efficiency. However, the maximum turbine inlet temperature is limited to the material characteristics of the turbine airfoils, especially the first stage guide vanes and rotor blades, since these airfoils are exposed to the highest temperature.
In order to allow for a higher gas flow temperature, the turbine airfoils include complex internal cooling circuits to provide the maximum amount of cooling for the airfoil while making use of the minimum amount of cooling air in order to maximize the efficiency of the turbine and therefore the engine. Internal airfoil cooling circuits have been proposed with complex design in order to maximize the amount of cooling as well as minimize the amount of cooling air used in order to increase the turbine efficiency and to increase turbine airfoil life. A serpentine flow cooling circuit is a very efficient arrangement to provide for cooling within the airfoils sine the serpentine path winds back and forth within the airfoil to increase the path length for the cooling air. FIG. 1 shows a prior art first stage turbine blade external heat transfer coefficient (HTC) profile. As shown in FIG. 1, the airfoil leading edge and trailing edge as well as the forward region of the suction side surface experience high hot gas heat transfer coefficient while the mid-chord section of the airfoil is at a lower hot gas HTC than the leading edge (LE) and the trailing edge (TE) and forward suction side (S/S) sections.
FIG. 2 shows a cross section view of a prior art turbine blade with a 5-pass aft flowing serpentine cooling circuit for the second stage blade. FIG. 3 shows a top view of a cross section through the turbine blade of FIG. 2, and FIG. 4 shows a diagram of the cooling air passage through the turbine blade of FIG. 2. The serpentine cooling circuit of FIG. 2 includes a first leg channel 11 extending along the leading edge region of the blade, a second leg forming a down pass channel 12, a third leg 13 forming an up-pass channel, a fourth leg 14 forming another down pass channel, and a fifth or last leg 15 extending along the trailing edge region of the blade. A first blade tip turn 16 and a second blade tip turn 18 turn the cooling air from an up-pass channel into the adjacent down-pass channel. A first blade root turn 17 and second blade root turn 19 turns the cooling air from a down-pass channel into the adjacent up-pass channel. A cover plate 21 covers over passages in the root to force the cooling air to follow the serpentine circuit. A row of exit cooling holes 22 discharge cooling air form the last leg 15 out from the airfoil cooling circuit. For an aft flowing 5-pass serpentine cooling circuit used for the entire airfoil, the cooling air flows through the serpentine cooling channels, lowering the airfoil metal temperature while increasing the cooling air temperature. As the cooling air reaches the airfoil trailing edge region, it loses some cooling capability (due to a pickup pf heat while passing through the airfoil mid-chord region) and therefore induces a hot spot for the airfoil trailing edge metal temperature. Hot spots appearing on a turbine airfoil especially in an industrial turbine engine induce problems with oxidation, which significantly reduces the part life in the engine. Also, it over-cools the airfoil mid-chord section where the heat loads for that region are low. In order to achieve a uniform sectional metal temperature distribution, a re-distribution of cooling air within the 5-pass serpentine flow circuit is required.
It is therefore an object of the present invention to provide for a turbine airfoil with a serpentine flow cooling circuit that cools less of the airfoil mid-chord region while cooling more of the trailing edge region than the cited prior art turbine blade serpentine flow cooling circuit.